Multi-source turbine cooling air

ABSTRACT

A gas turbine engine comprises a compressor section and a turbine section, with the turbine section having a first stage blade row and a downstream blade row. A higher pressure tap is tapped from a higher pressure first location in the compressor. A lower pressure tap is tapped from a lower pressure location in the compressor which is at a lower pressure than the first location. The higher pressure tap passes through a heat exchanger, and then is delivered to cool the first stage blade row in the turbine section. The lower pressure tap is delivered to at least partially cool the downstream blade row.

BACKGROUND OF THE INVENTION

This application relates to arrangements for cooling a gas turbineengine turbine section utilizing multiple sources for the cooling air.

Gas turbine engines are known and typically include a fan delivering airinto a bypass duct as propulsion air, and into a compressor as coreairflow. The air is compressed in the compressor and delivered into acombustion section where it is mixed with fuel and ignited. Products ofthis combustion pass downstream over turbine rotors driving them torotate. The turbine rotors, in turn, drive the fan and compressorrotors.

As can be appreciated, the turbine section experiences very hightemperatures. Thus, it is known to provide cooling air from otherlocations in the gas turbine engine to cool the turbine section.

It is known to tap cooling air from a downstream location in thecompressor section to the turbine sections. The downstream compressorair is hot itself. Thus, it is known to pass the cooling air through aheat exchanger on the way to cooling the turbine sections.

Particularly at upstream turbine rotor stages, this high pressurecompressed air is able to provide adequate cooling.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine comprises a compressorsection and a turbine section, with the turbine section having a firststage blade row and a downstream blade row. A higher pressure tap istapped from a higher pressure first location in the compressor. A lowerpressure tap is tapped from a lower pressure location in the compressorwhich is at a lower pressure than the first location. The higherpressure tap passes through a heat exchanger, and then is delivered tocool the first stage blade row in the turbine section. The lowerpressure tap is delivered to at least partially cool the downstreamblade row.

In another embodiment according to the previous embodiment, the higherpressure tap passes from the heat exchanger toward the turbine section,and is split into a first path heading radially outwardly to cool anupstream end of the first stage blade row, and a second path movingradially inwardly of a hub mounting the first stage blade row and thenmoves radially outwardly to cool a downstream end of the first stageblade row.

In another embodiment according to any of the previous embodiments,radially outwardly extending air from the higher pressure tap also coolsa vane mounted intermediate the first stage blade row and the downstreamblade row.

In another embodiment according to any of the previous embodiments, theradially outwardly extending air from the higher pressure tap also coolsan upstream end of the downstream blade row.

In another embodiment according to any of the previous embodiments, thelower pressure tap passes radially inwardly of the first stage bladerow, and axially beyond the downstream blade row and then radiallyoutwardly to cool a downstream end of the downstream stage blade row.

In another embodiment according to any of the previous embodiments, thedownstream stage blade row is a second stage, and the first stage bladerow and the second stage rotate together as a single rotor.

In another embodiment according to any of the previous embodiments, afan is positioned upstream of the compressor section and delivers airinto a bypass duct as propulsion air, and into the compressor sectionwith a bypass ratio defined as the volume ratio of air delivered intothe bypass duct compared to the volume of air delivered into thecompressor, with the bypass ratio being greater than or equal to about6.0.

In another embodiment according to any of the previous embodiments, thebypass ratio is greater than or equal to about 10.0.

In another embodiment according to any of the previous embodiments, afan drive turbine rotor is positioned downstream of a turbine rotorincluding the first stage blade row and the downstream blade row, withthe fan drive turbine driving the fan through a gear reduction.

In another embodiment according to any of the previous embodiments, agear ratio of the gear reduction is greater than or equal to about2.3:1.

In another embodiment according to any of the previous embodiments,radially outwardly extending air from the higher pressure tap also coolsa vane mounted intermediate the first stage blade row and the downstreamblade row.

In another embodiment according to any of the previous embodiments, theradially outwardly extending air from the higher pressure tap also coolsan upstream end of the downstream blade row.

In another embodiment according to any of the previous embodiments, thelower pressure tap passing radially inwardly of the first stage bladerow, and axially beyond the downstream blade row and then radiallyoutwardly to cool a downstream end of the downstream stage blade row.

In another embodiment according to any of the previous embodiments, thedownstream stage blade row is a second stage, and the first stage bladerow and the second stage rotate together as a single rotor.

In another embodiment according to any of the previous embodiments, thelower pressure tap passing radially inwardly of the first stage bladerow, and axially beyond the downstream blade row and then radiallyoutwardly to cool a downstream end of the downstream stage blade row.

In another embodiment according to any of the previous embodiments, thedownstream stage blade row is a second stage, and the first stage bladerow and the second stage rotate together as a single rotor.

In another embodiment according to any of the previous embodiments, afan is positioned upstream of the compressor section and delivers airinto a bypass duct as propulsion air, and into the compressor sectionwith a bypass ratio defined as the volume ratio of air delivered intothe bypass duct compared to the air delivered into the compressor, withthe bypass ratio being greater than or equal to about 6.0.

In another embodiment according to any of the previous embodiments, afan is positioned upstream of the compressor section and delivers airinto a bypass duct as propulsion air, and into the compressor sectionwith a bypass ratio defined as the volume ratio of air delivered intothe bypass duct compared to the air delivered into the compressor, withthe bypass ratio being greater than or equal to about 6.0.

In another embodiment according to any of the previous embodiments, thebypass ratio is greater than or equal to about 10.0.

In another embodiment according to any of the previous embodiments, afan drive turbine rotor is positioned downstream of a turbine rotorincluding the first stage blade row and the downstream blade row, withthe fan drive turbine driving the fan through a gear reduction.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows an embodiment of a gas turbine engine.

FIG. 2 shows an embodiment of a portion of a turbine section.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7 ° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 shows a gas turbine engine 80 incorporating a multi-sourcecooling scheme for a turbine section 93. A compressor 82 is shownschematically. A heat exchanger 84 receives a tap 86 of high pressurecompressor air. This may be tapped, as an example, from an exit of ahigh pressure compressor, such as in the engine 20 illustrated inFIG. 1. The air passes through the heat exchanger 84 and into a flowpassage 90 heading toward a high pressure turbine section 93.

A second tap 88 is tapped from a lower pressure location in thecompressor section 82. As an example, the tap 88 may be air atapproximately 1100° F. (593° C.). Thus, a heat exchanger may not benecessary for this air. The air is tapped in a path 89 also toward theturbine section 93. Air 90 cools an upstream end 92 of first stage bladerow 91 from a path 96. Another path 98 moves radially inward of a hub100 of the first stage and then radially outwardly at 102 and splits at104 to cool a downstream end 94 of the first blade row 91 and a vane106. Another branch 108 from the path 102 cools the vane 106 and anupstream end 112 of a second blade row 110.

The cooling path 89 extends radially outwardly as shown at 116 to coolthe downstream end 114 of the second stage blade row 110.

As should be understood, the air in path 90 is at a significantly higherpressure than air in path 89. This will facilitate cooling of the higherpressures seen by the first blade row 91, and even the upstream end 112of the second blade row 110. However, the lower pressures in flow path89 will be sufficient to move across the downstream end 114 of thesecond blade row 110, as products of combustion will be at a lowerpressure than at the upstream end 112.

In this manner, the air from the tap 86, which has already receivedsignificantly more work than the air from the tap 88, is used moreconservatively, thus, increasing the efficiency of the overall engineoperation. Since path 90 is cooled, and path 89 is not, the two areclose to the same temperature. This is beneficial to increase turbinedisk life.

The gas turbine 80, as shown in FIG. 2, is particularly valuable whenutilized with an engine as set forth in FIG. 1 wherein a fan driveturbine drives the fan rotor through a gear reduction. In such engines,the high ratio of air delivered into the bypass duct compared to thevolume of air delivered into the compressor section makes efficient useof the air delivered into the compressor section very important. Thus,the cooling arrangement disclosed in this application becomesparticularly valuable when utilized in such a gas turbine engine.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a compressor section and a turbinesection, with said turbine section having a first stage blade row and adownstream blade row; a higher pressure tap tapped from a higherpressure first location in said compressor; and a lower pressure taptapped from a lower pressure location in said compressor which is at alower pressure than said first location, said higher pressure tappassing through a heat exchanger, and then being delivered to cool saidfirst stage blade row in said turbine section, and said lower pressuretap being delivered to at least partially cool said downstream bladerow.
 2. The gas turbine engine as set forth in claim 1, wherein saidhigher pressure tap passing from said heat exchanger toward said turbinesection, and split into a first path heading radially outwardly to coolan upstream end of said first stage blade row, and a second path movingradially inwardly of a hub mounting said first stage blade row and thenmoving radially outwardly to cool a downstream end of said first stageblade row.
 3. The gas turbine engine as set forth in claim 2, whereinradially outwardly extending air from said higher pressure tap alsocooling a vane mounted intermediate said first stage blade row and saiddownstream blade row.
 4. The gas turbine engine as set forth in claim 4,wherein said radially outwardly extending air from said higher pressuretap also cooling an upstream end of said downstream blade row.
 5. Thegas turbine engine as set forth in claim 4, wherein said lower pressuretap passing radially inwardly of said first stage blade row, and axiallybeyond said downstream blade row and then radially outwardly to cool adownstream end of said downstream stage blade row.
 6. The gas turbineengine as set forth in claim 5, wherein said downstream stage blade rowis a second stage, and said first stage blade row and said second stagerotating together as a single rotor.
 7. The gas turbine engine as setforth in claim 6, wherein a fan is positioned upstream of saidcompressor section and said fan delivering air into a bypass duct aspropulsion air, and into said compressor section with a bypass ratiodefined as the volume ratio of air delivered into said bypass ductcompared to the volume of air delivered into said compressor, with saidbypass ratio being greater than or equal to about 6.0.
 8. The gasturbine engine as set forth in claim 7, wherein said bypass ratio isgreater than or equal to about 10.0.
 9. The gas turbine engine as setforth in claim 8, wherein a fan drive turbine rotor is positioneddownstream of a turbine rotor including said first stage blade row andsaid downstream blade row, with said fan drive turbine driving said fanthrough a gear reduction.
 10. The gas turbine engine as set forth inclaim 9, wherein a gear ratio of said gear reduction is greater than orequal to about 2.3:1.
 11. The gas turbine engine as set forth in claim1, wherein radially outwardly extending air from said higher pressuretap also cooling a vane mounted intermediate said first stage blade rowand said downstream blade row.
 12. The gas turbine engine as set forthin claim 11, wherein said radially outwardly extending air from saidhigher pressure tap also cooling an upstream end of said downstreamblade row.
 13. The gas turbine engine as set forth in claim 12, whereinsaid lower pressure tap passing radially inwardly of said first stageblade row, and axially beyond said downstream blade row and thenradially outwardly to cool a downstream end of said downstream stageblade row.
 14. The gas turbine engine as set forth in claim 13, whereinsaid downstream stage blade row is a second stage, and said first stageblade row and said second stage rotating together as a single rotor. 15.The gas turbine engine as set forth in claim 1, wherein said lowerpressure tap passing radially inwardly of said first stage blade row,and axially beyond said downstream blade row and then radially outwardlyto cool a downstream end of said downstream stage blade row.
 16. The gasturbine engine as set forth in claim 15, wherein said downstream stageblade row is a second stage, and said first stage blade row and saidsecond stage rotating together as a single rotor.
 17. The gas turbineengine as set forth in claim 16, wherein a fan is positioned upstream ofsaid compressor section and said fan delivering air into a bypass ductas propulsion air, and into said compressor section with a bypass ratiodefined as the volume ratio of air delivered into said bypass ductcompared to the air delivered into said compressor, with said bypassratio being greater than or equal to about 6.0.
 18. The gas turbineengine as set forth in claim 1, wherein a fan is positioned upstream ofsaid compressor section and said fan delivering air into a bypass ductas propulsion air, and into said compressor section with a bypass ratiodefined as the volume ratio of air delivered into said bypass ductcompared to the air delivered into said compressor, with said bypassratio being greater than or equal to about 6.0.
 19. The gas turbineengine as set forth in claim 18, wherein said bypass ratio is greaterthan or equal to about 10.0.
 20. The gas turbine engine as set forth inclaim 16, wherein a fan drive turbine rotor is positioned downstream ofa turbine rotor including said first stage blade row and said downstreamblade row, with said fan drive turbine driving said fan through a gearreduction.